This invention relates to a separation mechanism for small space vehicles, and, more particularly, to such a mechanism that operates from a single disengagement actuator device.
Space vehicles are often constructed as a launcher or booster stage that provides the primary propulsion, and a payload in a second stage that is carried by the launcher/booster. The two stages are mechanically and electrically joined together while the launcher/booster operates, and separated after the launcher/booster is no longer needed. This approach is used in a wide variety of types of space vehicle situations.
In an example, in one concept of a space defense system a small space vehicle payload is carried on the forward end of a booster that provides primary propulsion. The booster has a primary axial thruster and the required propulsion tanks. The small space vehicle has a sensor that seeks a target. When a target is detected, the booster thruster is fired to accelerate the combined vehicle toward the target. When the booster is no longer needed, the small space vehicle is separated from the booster and travels toward the target. The small space vehicle has guidance thrusters that permit it to change its trajectory to intercept the target, and these guidance thrusters are more effective for the reduced mass and inertia of the small space vehicle as compared with the combined small space vehicle and booster. After separation, the small space vehicle changes course as necessary to impact the target.
A clean separation of the small space vehicle from the booster is critical to the success of the system. Care is required so that the trajectory of the small space vehicle is not significantly altered by the act of separation, with the result that the sensor of the small space vehicle would lose alignment with the target either permanently or temporarily so that an expenditure of the limited fuel of the small space vehicle would be expended to place the small space vehicle back on course to the target. The separation must also be clean in the sense that the small space vehicle and its sensor are not damaged during the separation, as by debris or gases that are emitted from either of the stages during the separation.
Existing and proposed separation mechanisms have suffered shortcomings in some of these areas, and may also have insufficient reliability. Various steps of vehicle separation are typically accomplished by the firing of pyrotechnic devices or other actuators in a precisely defined order. Thus, for example, in one approach several mechanical linkages and the electrical connections between the stages are severed by the firing of a number of pyrotechnic devices. Another pyrotechnic device is fired to push the two stages apart. Thus, multiple pyrotechnic devices are utilized, and these multiple devices must be fired at a precisely defined time relative to the firing of the other devices or the separation will not be fully successful. Moreover, each pyrotechnic or other type of actuator device has a probability of successful operation upon command, its reliability, that is less than 1.0. Consequently, the more pyrotechnic devices that must be operated, the lower is the overall reliability of the separation mechanism. Each pyrotechnic device creates a structural dynamic shock and acoustic load upon activation. The more pyrotechnic devices that are used, the more shocks and loadings of the structure from this source.
There is a need for an improved approach to achieving separation of small space vehicles that does not significantly alter the trajectory of the vehicle of interest, optimizes system weight, has low joint compliance to minimize structural dynamic motion, does not cause damage in any manner to the vehicle, and is highly reliable. The present invention fulfills this need, and further provides related advantages.